1. Field of the Invention
The present application relates generally to flight control systems, and more specifically, to an aircraft flight control system for controlling rotor blade flapping introduced by collective pitch.
2. Description of Related Art
All rotor systems are subject to dissymmetry of lift in forward flight. During hover, the lift is equal across the entire rotor disk. As the helicopter gains airspeed, the advancing rotor blade develops greater lift because of the increased airspeed. For example, rotor blades at hover move at 300 knots and in forward flight at 100 knots the advancing blades move at a relative speed of 400 knots and while the retreating blades move at 200 knots. This has to be compensated for in some way, or the helicopter would corkscrew through the air doing faster and faster snap rolls as airspeed increased.
Dissymmetry of lift is compensated for by blade flapping. Because of the increased airspeed (and corresponding lift increase) on the advancing rotor blade, the rotor blade flaps upward. Decreasing speed and lift on the retreating rotor blade causes the blade to flap downward. This induced flow through the rotor system changes the angle of attack on the rotor blades and causes the upward-flapping advancing rotor blade to produce less lift, and the downward-flapping retreating rotor blade to produce a corresponding lift increase. Some rotor system designs require that flapping be limited by flapping stops which prevent damage to rotor system components by excessive flapping. In addition to structural damage, aircraft control can be compromised if the rotor flaps into the stop. Thus it becomes incumbent on the aircraft designer to control flapping and warn of this hazardous condition. This application addresses this requirement. Although the foregoing developments represent great strides in the area of flapping detection and reduction, many shortcomings remain.
Previous attempts to reduce flapping by limiting cyclic control inputs, such as was disclosed by U.S. Pat. No. 8,496,199, which is hereby incorporated by reference as if fully set forth, only considered rotor flapping and cyclic control positions as inputs. Furthermore, previous attempts have been forced to first measure flapping and then react to the flapping. For example, in forward flight at speeds greater than 40 KCAS in conversion mode, flapping due to collective can be as high as 1 degree per degree of collective pitch input. This flapping contribution can not be acted upon by previous CPMS implementations until it is sensed.
Equation (1) shows the upper limits of Control Power Management System (CPMS) CPMS-based longitudinal cyclic limits, respectively.BULIM=BBlong+√{square root over ((FMAX2−b12))}  (1)
Equation (2) shows the lower limits of Control Power Management System (CPMS) CPMS-based longitudinal cyclic limits, respectively.BLLIM=BBlong−√{square root over ((FMAX2−b12))}  (2)
where B_ULIM=upper CPMS-based longitudinal cyclic command limit, B_LLIM=lower CPMS-based longitudinal cyclic command limit, BB_long is the longitudinal component of blowback, F_max is the design maximum total flapping, and b_1 is the lateral component of flapping.
Experience with tiltrotors has shown that more effective flapping control is possible if collective pitch is made available to the CPMS.
While the system and method of the present application is susceptible to various modifications and alternative forms, specific embodiments thereof have been shown by way of example in the drawings and are herein described in detail. It should be understood, however, that the description herein of specific embodiments is not intended to limit the invention to the particular embodiment disclosed, but on the contrary, the intention is to cover all modifications, equivalents, and alternatives falling within the spirit and scope of the process of the present application as defined by the appended claims.